1. Field of the Invention
The present invention relates generally to an air cooled turbine airfoil, and more specifically to a turbine airfoil with near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, a high temperature gas flow is passed through a turbine to produce mechanical power to drive a bypass fan in the case of an aero engine or to drive a generator in the case of the industrial engine. The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the highest temperature attainable is dependant upon several factors such as the material properties of the turbine and the cooling ability of the airfoils.
The first stage turbine stator vanes and rotor blades are exposed to the highest gas flow temperature in the engine, and therefore require the most cooling. In the prior art, near wall cooling is used in the airfoil main body that have radial flow channels plus re-supply holes in series with film discharge cooling holes. FIG. 1 shows a prior art turbine blade and FIG. 2 shows a cross section of the internal cooling channels and film discharge holes. In the cooling circuit of FIG. 2, spanwise and chordwise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. Also, use of single radial channel flow is not the best method of utilizing cooling air since this results in a low convective cooling design.
It is an object of the present invention to provide for an air cooled turbine airfoil with a reduced airfoil main body metal temperature which results in reduced airfoil cooling flow requirement and improved turbine efficiency.